"The Government of the United States of America has rights in this invention pursuant to Contract No. DE-AC02-92CE40960 awarded by the U.S. Department of Energy."
In operation of a gas turbine engine, air at atmospheric pressure is initially compressed by a compressor and delivered to a combustion stage. In the combustion stage, heat is added to the air leaving the compressor by adding fuel to the air and burning it. The gas flow resulting from combustion of fuel in the combustion stage then expands through a turbine, delivering up some of its energy to drive the turbine and produce mechanical power.
The gases within the combustor typically range from between 2000 degrees to at least 2500 degrees Fahrenheit. Since the efficiency and work output of the turbine engine are related to the entry temperature of the incoming gases, at the turbine section, there is a trend in gas turbine engine technology to increase the gas temperature. A consequence of this is that the materials of which the combustor, blades and vanes are made assume ever-increasing importance with a view to resisting the effects of elevated temperature.
Combustors historically have been made of metals such as high temperature steels. More recently they are made of nickel alloys. It has been found necessary to provide internal cooling passages in order to prevent melting. Ceramic coating can enhance the heat resistance of the turbine components. In specialized applications, nozzle guide vanes and blades are being made entirely of ceramic, thus, imparting resistance to even higher gas entry temperatures and requiring higher temperatures within the combustor.
However, if the combustor is made of ceramic, which have a different chemical composition, physical property and coefficient of thermal expansion to that of a metal supporting structure, then undesirable thermal stresses will be set up between the combustor and its supports when the engine is operating. Such undesirable thermal stresses cannot adequately be contained by cooling.
Furthermore, conventional assembly techniques and methods will require alternative designs, processes and assembly techniques. The structural components of the combustor and the assembly of the combustor within the gas turbine engine will need to be rethought.
Historically, using metallic components, a combustor design has used a multipiece design of segments one overlaps another. The segments are rigidly secured one to another by braze, bolts and/or welding. As an alternative, the combustor has been formed from a single piece. With a ceramic combustor, the integrity of the material and the construction thereof can drastically increase cost and can result in premature failure due to flaws in the surface. The larger the physical size of the ceramic shape the lesser the likelihood of producing a component having structural integrity. The sliding friction between the ceramic combustor and the supporting structure creates a contact tensile stress on the ceramic that will cause surface deterioration. If this deterioration in the surface of the ceramic occurs in a tensile stress zone of the combustor the generated surface flaw can result in catastrophic failure.
The present invention is directed to overcoming one or more of the problems as set forth above.